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The development of guidance, navigation and control algorithms and avionic procedures requires the disposability of suitable analysis and verification tools, such as simulation environments, which support the design process and allow detecting potential problems prior to the flight test, in order to make new technologies available at reduced cost, time and risk. This paper presents a simulation environment for avionic software development and qualification, especially aimed at equipment for general aviation aircrafts and unmanned aerial systems. The simulation environment includes models for short and medium-range radio-navigation aids, flight assistance systems, and ground control stations. All the software modules are able to simulate the modeled systems both in fast-time and real-time tests, and were implemented following component oriented modeling techniques and requirement based approach. The paper describes the specific models features, the architectures of the implemented software systems and its validation process. Performed validation tests highlighted the capability of the simulation environment to guarantee in real-time the required functionalities and performance of the simulated avionics systems, as well as to reproduce the interaction between these systems, thus permitting a realistic and reliable simulation of a complete mission scenario., {"references":["E. Filippone, V. Di Vito, G. Torrano, D. Taurino, A. Ferreira, D. Zammit-Mangion, J. Gauci, G. Gargiulo, \"RPAS – ATM integration demonstration – real time simulation results\", 15th AIAA Aviation Technology, Integration, and Operations Conference, AIAA AVIATION Forum, (AIAA 2015-3407).","E. De Lellis, V. Di Vito, L. Garbarino, C. Lai, F. Corraro, \"Design process and real-time validation of an innovative autonomous mid-air flight and landing system\", World Academy of Science, Engineering and Technology N.55 , pp. 174 (2011).","A. Fedele, N. Genito, L. Garbarino, G. L. Di Capua, V. Baraniello, \"A pilot-in-the-loop facility for avionic concept development\", Journal of Automation and Control Engineering (Vol. 4, No. 4, August 2016).","N. Genito, F. Montemari, G. Corraro, D. Rispo, R. Palumbo, \"Integrated simulation facility for interoperability operation\", Lecture Notes in Computer Science LNCS-8906 (2014).","J. Kim, S. Lee, K. Ryu, \"Development of avionics hot bench for avionics system simulation and validation\", AIAA Modeling and Simulation Technologies Conference and Exhibit, AIAA 2007-6362.","Software Tool Qualification Considerations, RTCA, December 2011. DO-330.","A. Helfrick, L. Buckwalter, Principles of avionics, 8th ed., Avionics Communications Inc., Leesburg, VA, USA (2013).","Minimum Operational Performance Standards for Universal Access Transceiver (UAT) Automatic Dependent Surveillance — Broadcast, RTCA, Inc. December, 2009. DO-282B.","Minimum Operational Performance Standards for 1090 MHz Extended Squitter Automatic Dependent Surveillance — Broadcast (ADS-B) and Traffic Information Services — Broadcast (TIS-B). RTCA, Inc.. December, 2011. DO-260B.\n[10]\tM. Kayton, W. Fried, Avionics Navigation Systems, J. Wiley & Sons. Inc., London (UK), 1997.\n[11]\tICAO ANNEX 10 Aeronautical Telecommunications Volume I, Radio Navigation Aids, 6th Edition, 2006.\n[12]\tU.S. Department of Transportation, Federal Aviation Administration, Aeronautical Information Manual, Feb.2012.\n[13]\tG. B. Litchford, \"Analysis of cumulative errors of cat. II, III operations with requirements for additional research\", NASA CR-1188, 1968.\n[14]\tH. Johansen, \"A survey of general coverage NAVAIDS for V/STOL aircraft: a VOR/DME error model\", MIT, NASA-CR-1588, 1970.\n[15]\t2010 Federal Radionavigation Plan, Department of Defense, Department of Homeland Security, and Department of Transportation, Apr.2011.\n[16]\tICAO, Manual on Testing of Radio Navigation Aids, Volume I, Doc 8071, Fourth Edition, 2000.\n[17]\tU.S. Department of Transportation, Federal Aviation Administration, U.S. National Aviation Handbook for the VOR/DME/TACAN Systems, FAA Order 9840.1, 1982.\n[18]\tENAV, AIP Italia, Jan. 2014.\n[19]\tR. Austin, Unmanned aircraft systems: UAVs design, development and deployment, Wiley, 2010.\n[20]\tJ. Gundlach, Designing unmanned aircraft systems: a comprehensive approach, AIAA Education Series, 2012.\n[21]\tRecommendation ITU-R P.676-9 (02/2012), \"Attenuation by atmospheric gases\".\n[22]\tRecommendation ITU-R P.840-5 (02/2012), \"Attenuation due to clouds and fog\".\n[23]\tR. E. Collin, Antennas and Radiowave Propagation, Mc Graw Hill, 1985.\n[24]\tJ. Ziccardi et al., \"Measuring UAS pilot responses to common air traffic clearances\", Lecture Notes in Computer Science Volume 8017, 2013, pp 606-612.\n[25]\tG. Carrigan, D. Long, M. L. Cummings, J. Duffner, \"Human factors analysis of predator B crash\", AUVSI 2008.\n[26]\tSoftware Considerations in Airborne Systems and Equipment Certification, RTCA, December 2011. DO-178C.\n[27]\tB. Potter, \"Complying with DO-178C and DO-331 using Model-Based Design\", 12AEAS-0090 (2012).\n[28]\tR. G. Estrada, E. Dillaber, G. Sasaki, Best practices for developing DO-178 compliant software using Model-Based Design, 2013.\n[29]\tM. Orefice, V. Di Vito, L. Garbarino, F. Corraro, G. Fasano, D. Accardo, \"Real-time validation of an ADS-B based aircraft conflict detection system\", AIAA Infotech @ Aerospace, AIAA SciTech Forum, (AIAA 2015-0486)."]}

The ionosphere is commonly divided into the portion below (bottomside) and above (topside) the region at which peak values of electron density occur. Topside ionospheric modeling is a challenging problem because of the limited data available. Indeed, the more intense peak ionization region, or bottomside ionosphere, dominates the effects observable from ground stations. High-altitude ionosondes, such as sounding rockets, have been traditionally used for direct sounding only of the higher ionospheric layers. Nowadays, signals of opportunity exist for sounding the ionosphere with no dedicated ionosondes. With the continuous deployment of GPS receivers on board spacecraft for positioning, indirect sounding of the topside ionosphere using navigation signals can be performed. This paper reviews geometric-based models allowing to infer the total electron content of the topside ionosphere from spacecraft GPS measurements.

The paper investigates different solutions for handling of ionospheric delays in high accuracy long baseline relative positioning by Carrier-based Differential GPS (CDGPS). Standard literature approaches are reviewed and the relevant limitations discussed. Hence, a completely ionospheric-free approach is proposed, in which the differential ionospheric delays are cancelled out by combination of dual frequency GPS measurements. The performance of this approach is quantified over real-world spaceborne GPS data made available by the Gravity Recovery And Climate Experiment (GRACE) mission and compared to the standard solutions.

This paper deals with the real-time onboard accurate relative positioning by Carrier-phase Differential GPS (CDGPS) of LEO formations with baselines of hundreds of kilometers. On long baselines, high accuracy can be achieved only using dual-frequency measurements and exploiting the integer nature of Double Difference (DD) carrier-phase ambiguities. However, large differential ionospheric delays and broadcast ephemeris errors complicate the integer resolution task. The approach presented in this paper relies on separating the integer ambiguities’ resolution from the relative positioning solution. The first task is performed by a closed-loop dynamic-based filter integrating an Extended Kalman Filter with an Integer Least Square (ILS) estimator. Then, the relative position is computed with a conventional kinematic least-square algorithm processing ionospheric-free combination of carrier-phase measurements de-biased of the integer ambiguities. Three different formulations of the dynamic-based filter are compared. The first two formulations use Lear’s model to estimate differential ionospheric delays, the main difference being in the validation step performed on all the ambiguities, in the first case, and only on the wide-lane combination, in the second case. The third scheme is instead based on combining the DD measurements for removing ionospheric delays from the observation model. The performance of the developed solutions is assessed by using flight data from the Gravity Recovery and Climate Experiment mission. Results show that the closed-loop approach based on validating only the wide-lane combinations shows the best performance.

The invention concems a method and system for the prediction of aircrafts vertical trajectory, in particular for Air Traffic Management, comprising the following flight calculation modules: Take-off; Climb; Cruise; Descent; and Landing, corresponding to the relevant flight phases, wherein: the calculation of the predicted aircraft trajectory is effected by using a set ofTEM equations comprising, as output variables, the vertical rate of climb or descent, the true air speed, the energy share factor, the thrust and the drag, the mass ofthe aircraft modeled as point-mass, and comprising, as inpu variables, the Mach number depending on true air speed and temperature and altitude, the gravity acceleration, and the fuel flow, and the flight path angle; the calculation of the predicted aircraft trajectory for Cruise phase, wherein only the mass is varying, is performed by using the following analytical solution to said set of TEM equations.

The activities presented in this chapter are aimed at providing an approach for quantitatively evaluating, within a reasonably reduced timeframe, the conservatism of the safety margins in approach and departure operations. The ultimate goal of the proposed approach is to enable eventually more accurate risk-assessment analysis with a higher level of confidence that safety margins can be reduced, and thus that the significant effort required for assessing the risk will not be lost. The proposed methodology relaxes the standard risk-assessment problem up to the point in which simulation-based methodologies, most notably Monte Carlo analysis, can be used for the risk assessment. The approach feasibility is evaluated by devising a fictitious benchmark application case, concerning simultaneous operations of intersecting runways for departing and arriving flights. In this benchmark case, the Monte Carlo technique has been applied to a limited number of critical situations in which some hazardous events have occurred. Three mathematical models have been also developed for describing the aircraft relative dynamics and timing in these scenarios, which integrate various models proposed in the open literature for simulating arrival and departure operations. Application on the benchmark case demonstrates the ability of the proposed approach in providing quantitative information on the conservativeness of actual safety margins, and thus to drive refinement strategies of novel, more efficient, procedures which can be potentially compliant with a required Target Level of Safety

This paper presents a model of an airport apron that has been developed in order to determine a dynamical model for aircraft travel time in this area. This model is intended for future systems that will support automated routing and taxiing of aircraft in airport taxiways. The apron model proposed keeps into account the main issues related to the aircraft ground movements, such as aircraft conflict management. This model was developed to integrate the software tool taXi Route Planner (XRP), a proprietary software of Selex Sistemi Integrati S.p.A. Malpensa Airport was selected as a case study and the model was verified by field testing on real traffic data.

This paper focuses on the autonomous relative navigation of LEO satellite formations. Specifically, a CDGPS technique is addressed using a cascade-combination of dynamic and kinematic filters which processes double differenced code and carrier measurements on two frequencies to get centimeter-level baseline estimates. In order to improve navigation filter robustness against poor GPS geometry and noisy measurements, the CDGPS filter is augmented by inter-satellite local ranging measurements with decimeter-level range accuracy, as the ones provided by ranging transponders or GNSS-like systems. Performance of the augmented filter is tested by numerical simulations of orbital configurations characterized by short, varying baselines, relevant to next generation Synthetic Aperture Radar missions. Numerical results show that the augmented filter is able to preserve the centimeter-level accuracy of the relative navigation also in the cases in which the CDGPS-only filter exhibits a degraded accuracy due to a poor GDOP or a limited number of GPS satellites in view

In this paper performance of Carrier-phase Differential GPS (CDGPS) for relative navigation in applications where the inter-satellite distance highly varies along the orbit is investigated. Specifically, to satisfy high accuracy requirements, a hybrid filtering approach is proposed in which a dynamic filter is combined in cascade with a kinematic one. For the dynamic filter both an Extended Kalman Filter (EKF) and an Unscented Kalman Filter (UKF) are evaluated and compared to output high-accuracy estimates of the floating values of the carrier-phase integer ambiguities. These, once extracted with the LAMBDA method, are processed in the kinematic filer to come to highly precise relative position and velocity estimates.

A novel robustness analysis technique is proposed for atmospheric re-entry applications. The problem is stated as a finite time stability (FTS) analysis of linear time-varying (LTV) systems on a compact time domain, subject to bounded variations in initial state and unknown parameters. The FTS property is formulated as the inclusion of all the possible system trajectories into a pre-specified time-varying subset of the state space. Based on assuming the involved sets are polytopes, the proposed approach allows deducing the system FTS from the property verification on a limited number of numerically computed system trajectories. An additional result is presented which allows determination of a conservative estimate of the maximum norm-bound of time-varying perturbations under which the LTV system remains finite time stable. Results of the application of the proposed technique to a re-entry technology demonstrator are presented which demonstrate its effectiveness in complementing conventional linear time invariant-based analyses. Results also show that it is computationally viable and allows linking the system robustness to a quantitative analysis of the system trajectory dispersion around the nominal one due to concurrent initial state dispersion and uncertain parameters effects, which aids in evaluating mission objectives fulfilment.

GPS relative navigation filters could benefit notably from an accurate modeling of the ionospheric delays, especially over large baselines (>100 km) where double difference delays can be higher than several carrier wavelengths. This paper analyzes the capability of ionospheric path delay models proposed for spaceborne GPS receivers in predicting both zero-difference and double difference ionospheric delays. We specifically refer to relatively simple ionospheric models, which are suitable for real-time filtering schemes. Specifically, two ionospheric delay models are evaluated, one assuming an isotropic electron density and the other considering the effect on the electron density of the Sun aspect angle. The prediction capability of these models is investigated by comparing predicted ionospheric delays with measured ones on real flight data from the Gravity Recovery and Climate Experiment mission, in which two satellites fly separated of more than 200 km. Results demonstrate that both models exhibit a correlation in the excess of 80% between predicted and measured double-difference ionospheric delays. Despite its higher simplicity, the isotropic model performs better than the model including the Sun effect, being able to predict double differenced delays with accuracy smaller than the carrier wavelength in most cases. The model is thus fit for supporting integer ambiguity fixing in real-time filters for relative navigation over large baselines. Concerning zero-difference ionospheric delays, results demonstrate that delays predicted by the isotropic model are highly correlated (around 90%) with those estimated using GPS measurements. However, the difference between predicted and measured delays has a root mean square error in the excess of 30 cm. Thus, the zero-difference ionospheric delays model is not likely to be an alternative to methods exploiting carrier-phase observables for cancelling out the ionosphere contribution in single-frequency absolute navigation filters.

The unmanned space vehicle (USV) program has been undertaken by the Italian Center for Aerospace Research with the aim of developing flying test beds of next generation reentry launch vehicles. In this framework, the development of small demonstrators is also foreseen to validate technological and operational aspects of full-scale vehicles and missions. In this paper, a small-scale demonstrator of the sub-orbital re-entry test mission of the USV program is described. Both mission profile and objectives are very challenging in terms of demonstrator guidance, navigation and control. After a short description of the mission and demonstrator architectures, particular emphasis is given to the guidance and navigation analysis. To this end, mission objectives and reduced-scale system constaints are integrated and translated into innovative guidance solutions relying on optimization techniques. Then, performance of a commercial-off-the-shelf GPS-aided, miniature inertial navigation system over the proposed trajectories is evaluated by Monte Carlo analysis. Standalone inertial and GPS-aided inertial navigation performance is also compared considering GPS loss conditions due to antenna plasma effects.

The paper investigates different solutions for ionospheric delay handling in high accuracy long baseline relative positioning by Carrier-Phase Differential GPS (CDGPS). Standard literature approaches are reviewed and the relevant limitations are discussed. Hence, a completely ionosphere-free approach is proposed, in which the differential ionospheric delays are cancelled out by combination of dual frequency GPS measurements. The performance of this approach is quantified over real-world spaceborne GPS data made available by the Gravity Recovery and Climate Experiment (GRACE) mission and compared to the standard solution.

This paper deals with the relative navigation of a formation of two spacecrafts separated by hundreds of kilometers based on processing dual-frequency differential carrier-phase GPS measurements. Specific requirements of the considered application are high relative positioning accuracy and real-time on board implementation. These can be conflicting requirements. Indeed, if on one hand high accuracy can be achieved by exploiting the integer nature of double-difference carrier-phase ambiguities, on the other hand the presence of large ephemeris errors and differential ionospheric delays makes the integer ambiguities determination challenging. Closed-loop schemes, which update the relative position estimates of a dynamic filter with feedback from integer ambiguities fixing algorithms, are customarily employed in these cases. This paper further elaborates such approaches, proposing novel closed loop techniques aimed at overcoming some of the limitations of traditional algorithms. They extend techniques developed for spaceborne long baseline relative positioning by making use of an on-the-fly ambiguity resolution technique especially developed for the applications of interest. Such techniques blend together ionospheric delay compensation techniques, nonlinear models of relative spacecraft dynamics, and partial integer validation techniques. The approaches are validated using flight data from the Gravity Recovery and Climate Experiment (GRACE) mission. Performance is compared to that of the traditional closed-loop scheme analyzing the capability of each scheme to maximize the percentage of correctly fixed integer ambiguities as well as the relative positioning accuracy. Results show that the proposed approach substantially improves performance of the traditional approaches. More specifically, centimeter-level root-mean square relative positioning is feasible for spacecraft separations of more than 260 km, and an integer ambiguity fixing performance as high as 98% is achieved in a 1-day long dataset. Results also show that approaches exploiting ionospheric delay models are more robust and precise of approaches relying on ionospheric-delay removal techniques.

This paper presents a technique for on-the-fly rejection of GPS data outliers. It is particularly suited to applications where a high accuracy navigation solution is needed in real time, as for the relative positioning of satellites in Low Earth Orbit. The proposed technique relies on two metrics which screen the residual ranging errors in the Zero-Difference and Single-Difference pseudorange measurements with respect to a prefixed threshold. Preliminary tests carried out on flight data from the Gravity Recovery and Climate Experiment mission demonstrate the effectiveness of the proposed approach.

This paper deals with the problem of real-time onboard relative positioning of low-Earth-orbit spacecraft over long baselines using the Global Positioning System. Large intersatellite separations, up to hundreds of kilometers, are of interest to multistatic and bistatic synthetic-aperture radar applications, in which highly accurate relative positioning may be required in spite of the long baseline. To compute the baseline with high accuracy, the integer nature of dualfrequency, double-difference carrier-phase ambiguities can be exploited. However, the large intersatellite separation complicates the integer-ambiguities determination task due to the presence of significant differential ionospheric delays and broadcast ephemeris errors. To overcome this problem, an original approach is proposed, combining an extended Kalman filter with an integer least-square estimator in a closed-loop scheme, capable of fast on-the-fly integer-ambiguities resolution. These integer solutions are then used to compute the relative positions with a singleepoch kinematic least-square algorithm that processes ionospheric-free combinations of debiased carrier-phase measurements. Approach performance and robustness are assessed by using the flight data of the Gravity Recovery and Climate Experiment mission. Results show that the baseline can be computed in real time with decimeter-level accuracy in different operating conditions.

During the last 40 years, the mass of the artificial objects in orbit increased quite steadily at the rate of about 145 metric tons annually, leading to about 7000 metric tons. Most of the cross-sectional area and mass (97% in low Earth orbit) is concentrated in about 4500 intact abandoned objects plus a further 1000 operational spacecraft. Analyses have shown that the most effective mitigation strategy should focus on the disposal of objects with larger cross-sectional area and mass from densely populated orbits. Recent NASA results have shown that the worldwide adoption of mitigation measures in conjunction with active yearly removal of approximately 0.2−0.5% of the abandoned objects would stabilize the debris population. Targets would have typical masses between 500 and 1000 kg in the case of spacecraft, and of more than 1000 kg for rocket upper stages. In the case of Cosmos-3M second stages, more than one object is located nearly in the same orbital plane. This provides the opportunity of multi-removal missions, more suitable for yearly removal rate and cost reduction needs. This paper deals with the feasibility study of a mission for the active removal of large abandoned objects in low Earth orbit. In particular, a mission is studied in which the removal of two Cosmos-3M second stages, that are numerous in low Earth orbit, is considered. The removal system relies on a Chaser spacecraft which performs rendezvous maneuvers with the two targets. The first Cosmos-3M stage is captured and an autonomous de-orbiting kit, carried by the Chaser, is attached to it. The de-orbiting kit consists of a Hybrid Propulsion Module, which is ignited to perform stage disposal and controlled reentry after Chaser separation. Then, the second Cosmos-3M stage is captured and, in this case, the primary propulsion system of the Chaser is used for the disposal of the mated configuration. Critical mission aspects and related technologies are investigated at a preliminary level. In particular, an innovative electro-adhesive system for target capture, mechanical systems for the hard docking with the target and a hybrid propulsion technology suitable for rendezvous, de-orbiting and controlled reentry operations are analyzed. This is performed on the basis of a preliminary mission profile, in which suitable rendezvous and disposal strategies have been considered and investigated by numerical analysis. A preliminary system mass budget is also performed, showing that the Chaser overall mass is about 1350 kg, including a primary propulsion system of about 300 kg, and a de-orbiting kit with a mass of about 200 kg. The system designed results suitable to be launched with VEGA, actually the cheapest European space launcher.

Carrier-phase differential GPS (CDGPS) is a promising technology for accurate relative navigation in LEO formations of cooperating satellites, but navigation filter robustness against poor GPS geometry and noisy measurements has to be improved. This can be performed by augmenting the navigation filter with intersatellite local ranging measurements, as the ones provided by ranging transponders or GNSS-like systems. In this paper, an augmented CDGPS navigation filter is proposed for the formation of two satellites characterized by a short, varying baseline, relevant to next generation Synthetic Aperture Radar missions. Specifically, a cascade-combination of dynamic and kinematic filters which processes double-differenced code and carrier measurements on two frequencies, as well as local inter-satellite ranging measurements, is used to get centimeter-level baseline estimates. The augmented filter is validated by numerical simulations of the formation orbital path. Results demonstrate that the proposed approach is effective in preserving the centimeter-level accuracy achievable by a CDGPS-only filter also in the presence of a poor GDOP or a limited number of GPS satellites in view.

This paper describes a carrier-phase differential GPS approach for real-time relative navigation of LEO satellites flying in formation with large separations. These applications are characterized indeed by a highly varying number of GPS satellites in common view and large ionospheric differential errors, which significantly impact relative navigation performance and robustness. To achieve high relative positioning accuracy a navigation algorithm is proposed which processes double-difference code and carrier measurements on two frequencies, to fully exploit the integer nature of the related ambiguities. Specifically, a closed-loop scheme is proposed in which fixed estimates of the baseline and integer ambiguities produced by means of a partial integer fixing step are fed back to an Extended Kalman Filter for improving the float estimate at successive time instants. The approach also benefits from the inclusion in the filter state of the differential ionospheric delay in terms of the Vertical Total Electron Content of each satellite. The navigation algorithm performance is tested on actual flight data from GRACE mission. Results demonstrate the effectiveness of the proposed approach in managing integer unknowns in conjunction with Extended Kalman Filtering, and that centimeter-level accuracy can be achieved in real-time also with large separations.

The feasibility of a mission for the active removal of a large massive debris in Low Earth Orbit region is investigated. Critical aspects and related technologies are investigated at preliminary level. A preliminary mass budget shows that the chaser spacecraft overall mass varies in the range 1500-1700 kg, with Hybrid Propulsion Module of about 900 kg.

This paper focuses on a hardware-in-the-loop facility aimed at real-time testing of architectures and algorithms of multisensor sense and avoid systems. It was developed within a research project aimed at flight demonstration of autonomous non-cooperative collision avoidance for Unmanned Aircraft Systems. In this framework, an optionally piloted Very Light Aircraft was used as experimental platform. The flight system is based on multiple-sensor data integration and it includes a Ka-band radar, four electro-optical sensors, and two dedicated processing units. The laboratory test system was developed with the primary aim of prototype validation before multi-sensor tracking and collision avoidance flight tests. System concept, hardware/software components, and operating modes are described in the paper. The facility has been built with a modular approach including both flight hardware and simulated systems and can work on the basis of experimentally tested or synthetically generated scenarios. Indeed, hybrid operating modes are also foreseen which enable performance assessment also in the case of alternative sensing architectures and flight scenarios that are hardly reproducible during flight tests. Real-time multisensor tracking results based on flight data are reported, which demonstrate reliability of the laboratory simulation while also showing the effectiveness of radar/electro-optical fusion in a non-cooperative collision avoidance architecture.

This paper deals with the GPS-based relative navigation of LEO formations. Specifically, we consider applications characterized by two co-flying satellites with a large and highly variable separation, which are relevant to next generation monostatic/bistatic Synthetic Aperture Radar missions. In these applications, both scientific goals and control needs require the determination of the relative state with high accuracy. To this end, an Extended Kalman Filter is developed that processes double-difference pseudorange and carrier phase measurements on L1 and L2 frequencies. To preserve accuracy and robustness of the integer solution problem against large variations of the baseline, an original approach is developed in which, for each receiver, the Vertical Total Electron Content is included in the filter state. In addition, the double-difference ambiguities are re-estimated by the filter at each time step. A major technical problem of a filter processing double-differences is re-organizing the filter state when the pivot satellite changes. This is solved by an original and effective procedure that speeds up the filter convergence. Once the floating point estimates of the double-difference ambiguities have been produced by the dynamic filter, their integer values are extracted with the Least-Square Ambiguity Decorrelation Adjustment method and processed within a kinematic filter to estimate the relative position with high accuracy. Filter robustness and performance are evaluated by means of Monte Carlo simulations performed on the reference orbital scenario identified within the Italian SABRINA mission study. Results show that the integer ambiguities are always resolved, allowing to achieve a centimeter-level accuracy in all the simulated conditions.

A NOVEL approach to analyze the robustness of a flight control system (FCS) with respect to parametric uncertainties is presented, which specifically applies to gliding vehicles in the terminal phases of reentry flight. Robustness analyses are particularly challenging for these systems. Their reference trajectories are appreciably time-varying and encompass a broad variety of flight regimes. Furthermore, significant uncertainties on some critical design parameters affect the vehicle model, most notably those related to the aerodynamic behavior [1]. Current practice in FCS robustness analysis for this kind of application mainly relies on the theory of linear time-invariant (LTI) systems. In this approach, the original nonlinear system is linearized around a limited number of representative time-varying trajectories, including the nominal one. Then thewell-known frozentime approach [2] is applied, yielding multiple LTI models. In this way, classical stability margins [3] or more sophisticated LTI-based robustness criteria, such as analysis [4] and D-stability analyses [5], can be evaluated. Recently, a Lyapunov-based criterion coupled to interval analysis techniques [6] has been proposed for establishing robustness of a FCS. This approach does not resort to linearization of the system dynamics, but still requires the introduction of fictitious equilibrium points obtained by a frozen-time approach. Even if the flight experience demonstrated that frozen-time approaches are indeed operative, they are widely recognized as inefficient [7]. In fact, because the nominal trajectory may not be an equilibrium trajectory for the system in offnominal conditions, frozen-time analyses can provide only indicative, and often heavily conservative, results. To overcome such problems, further investigations are usually performed to identify a limited set of worst-case combinations of uncertain parameters to be used for FCS design refinement. In this case, nonlinear simulations in specific offnominal conditions, selected using sensitivity analysis and designer’s experience, represent the current practice. Optimization-based worst-case search has also been proposed [8], which may disclose the mutual effects of multiple uncertainties, but to a limited extent. In fact, the complexity of reentry dynamics under multiple uncertainties implies that actual worst cases relevant for FCS design refinement are difficult to identify. In any case, worst-case analysis can select only a limited number of test cases, hiding possible further causes of requirement violations, thus driving wrong refinement strategies that would not solve (or even worsen) FCS robustness problems. Monte Carlo (MC) analysis is, in practice, the only tool that is capable of investigating the combined effect of all uncertainties with a reasonable effort. However, being only a verification tool, when unsatisfactory robustness is discovered at this stage, the identification of its causes can require considerable postprocessing effort [9]. This yields one of the major limitations of this approach: that is, the limited support to the FCS design refinement when a requirement violation occurs due to poor robustness. As a result, in these cases, one is forced to iterate the design with scarce additional information. The present paper contributes toward advancing the current practice used in robustness analysis for FCS design refinement by introducing a method that takes into account nonlinear effects of multiple uncertainties over the whole trajectory, to be used before robustness is finally assessed withMC analysis. Themethod delivers feedback on the causes of requirement violation and adopts robustness criteria directly linked to the original mission or system requirements, such as those employed in MC analyses. The first objective is achieved estimating the region of requirement compliance in the space of the uncertain parameters. In this way, the approach provides an exhaustive coverage of the uncertainty’s effects on the FCS robustness. To translate mission requirements into robustness criteria over the whole trajectory, rather than at isolated points as in frozen-time approaches, we make use of the practical stability concept [10], which, to the authors’ knowledge, has never been applied to robustness analyses of atmospheric reentry vehicles.

A novel approach to flight control laws robustness analysis is presented, aimed at improving the efficiency of the clearance process with respect to conventional techniques, especially when applied to vehicles following time–varying reference trajectories. The approach exploits the convexity of LTV systems, coupled to an approximate description of the original nonlinear system by a certain number of its time-varying linearizations. The method effectiveness and potentials are ascertained by application to the longitudinal flight control laws clearance of the Italian Aerospace Research Center (CIRA) experimental vehicle USV.

A mission profile for advanced thermal protection system suborbital flight testing is identified. Its main goal is to achieve a constant heat flux at a specific area of the vehicle for a limited amount of time. A tool capable of exploring broad regions of the design space for these missions is developed, aiming at reducing possible design options to an extent manageable by conventional, more accurate, numeric-simulation-based methods. Based on a simplified model of the point mass dynamics, trajectories optimal for thermal protection system testing and compliant with prefixed path constraints are identified. The approximate method is validated comparing the obtained optimal trajectories with numeric-optimized standard solutions on three test cases. Then, to demonstrate the method effectiveness and flexibility, the mission design space is investigated for reasonable ranges of relevant parameters. Results show that increasing the vehicle’s ballistic coefficient allows reducing the specific mechanical energy at reentry, and that the maximum admissible dynamic pressure plays a principal role in affecting the attainable testing performances. An illustrative mission design for novel ceramic thermal protection system testing is presented that minimizes in the analyzed design space the specific mechanical energy at the trajectory apogee.

In the paper the maneuverability and stability (static and dynamic) characteristics of the USV Flying Test Bed (FTB_1) vehicle are presented. The analyses have been focused on the first dropped mission that achieves transonic flight regimes. Peculiarity of the here reported analyses is their application to the whole uncertainty space of the aerodatabase. Aerostructural datasets analyzed and relative uncertainty ranges, came from both CFD computations and from wind tunnel testing activities. Both classical and widely applied maneuverability and stability criteria, based on stability derivatives analyses, and root loci analyses for longitudinal and lateral-directional modes have been applied and results compared, giving suggestion for applicability of simplified analysis criteria.

A key aspect in the success of a space project is the capability to detect as soon as possible the problems that can arise during the project development. This approach allows to optimize mission reliability, project costs and temporal delays. As an example, given the extended range of flight regimes experienced by new generation Reusable Launch Vehicles ( RLV ) demonstrators (as NASA’s X-40A, X-43 ) throughout the various mission phases, assessing the impact of the aerodynamic uncertainties on the overall system performance is of great importance. System design should be performed so that uncertainties, with particular concern to the aerodynamics ones, do not significantly affect some basic vehicle properties, such as trajectory trimmability and dynamic stability. Therefore, identifying such admissible ranges of uncertainties might be a powerful system analysis methodology which could effectively help aerodynamics database refinement and system configuration development. A methodology aimed at quantifying the admissible ranges of uncertainties in which relevant vehicle properties can be guaranteed is presented. Specifically, the properties to which we refer are Trimmability and D-Stability of the longitudinal dynamics of an RLV-shaped, un-piloted, un-powered aircraft. The latter is a property wider than simple stability, in order to take into account that vehicle instability is acceptable if the Stability Augmentation System can still enforce the desired dynamics. The approach basically reduces the problem of determining the dynamic characteristics of the complete nonlinear system to an analysis of the robust stability of linear systems subject to uncertain parameters, by means of system’s linearization around a predetermined set of flight conditions. An analytic modeling of the system dynamics is carried out applying the well known Short Period Approximation, and successively deriving the necessary conditions that enforce stability and D-stability conditions as explicit functions of parameters of interest. An application case on a winged, autonomous RLV demonstrator vehicle is analyzed, modeling uncertainties on the vehicle longitudinal aerodynamic stability derivatives. Results have shown the method’s ability to identify the maximum admissible uncertainties, and to address the areas of major concern. Specifically, maximum acceptable uncertainties in terms of CLα and CMα have been identified, and can help to identify the needed level of accuracy in the aerodynamics model. Furthermore, when open-loop stability results are compared to closed-loop ones, indications can be derived on necessary flight control system features, in terms of both robustness and capability to cope with unstable plants.

This paper deals with the real-time onboard accurate relative positioning by carrier-phase differential GPS (CDGPS) of LEO formations with baselines of hundreds of kilometres. On long baselines, high accuracy can be achieved only using dual-frequency measurements and exploiting the integer nature of double difference (DD) carrier-phase ambiguities. However, large differential ionospheric delays and broadcast ephemeris errors complicate the integer resolution task. This paper is concerned with analysing possible approaches to DD ionospheric delays compensation in such applications. The first formulation models differential ionospheric delays as a function of the vertical electron content above the receivers, whereas the second one is based on combining DD measurements for removing ionospheric delays from the observation model. The effectiveness of the developed solutions is assessed by comparing the relative positioning accuracy that can be obtained on flight data. Results show that modelling the delays is advantageous for relative positioning only in mild ionospheric conditions.

The feasibility study of an active removal mission of multiple large abandoned objects from Low Earth Orbit region is presented. A chaser spacecraft, powered by a hybrid rocket engine, achieves contact with two Cosmos-3M second stages and provides for their de-orbiting by means of a hybrid propulsion module. Critical mission aspects and related technologies are investigated, such as an innovative electro-adhesive system for target capture, mechanical systems for the hard docking before the disposal. De-orbiting trajectory simulations and preliminary sizing and mass budget of chaser spacecraft and hybrid rocket engines were performed.

This paper presents and approach for real-time onboard relative positioning of two spacecraft separated of hundreds of kilometers with centimeter-level accuracy. The approach extends techniques developed for long baseline terrestrial relative positioning to spaceborne applications, making use of an on-the-fly ambiguity resolution technique especially developed for the applications of interest. Such technique blends together parametric models of the ionospheric delay, nonlinear models of relative spacecraft dynamics, and partial integer validation techniques, while being capable to comply with typical requirements for real-time on board operations. The approach is validated on flight data from the Gravity Recovery and Climate Experiment (GRACE) mission. Centimeter-level root-meansquare relative positioning is proved to be feasible for spacecraft separations exceeding 260 kilometers. Results show that in a one-day-long dataset 98% of integer ambiguities are fixed almost instantaneously, and less than 2% of these ambiguities are erroneous.

A novel approach to flight control laws robustness analysis is presented, aimed at improving the efficiency of the clearance process with respect to conventional techniques, especially when applied to vehicles following time–varying reference trajectories. The approach exploits the convexity of LTV systems, coupled to an approximate description of the original nonlinear system by a certain number of its time-varying linearizations. The method effectiveness and potentials are ascertained by application to the longitudinal flight control laws clearance of the Italian Aerospace Research Center (CIRA) experimental vehicle USV.

A key aspect for the success of a space project is the capability to detect as soon as possible the problems that can arise during the project development. This approach allows optimizing mission reliability, project costs and temporal delays. As an example, given the extended range of flight regimes experienced by new generation Reusable Launch Vehicles ( RLV ) demonstrators (as NASA’s X-40A, X-43 ) throughout the various mission phases, assessing the impact of the aerodynamic uncertainties on the overall system performance is of great importance. System design should be performed so that uncertainties, with particular concern to the aerodynamics ones, do not significantly affect some basic vehicle properties, such as trajectory trimmability and dynamic stability. Therefore, identifying such admissible ranges of uncertainties might be a powerful system analysis methodology which could effectively help to save costs for aerodynamics database and system configuration development. A methodology aimed at quantifying the admissible ranges of uncertainties in which some basic vehicle properties can be guaranteed is presented. Specifically, the properties we refer to are stability and D-Stability of the lateral-directional dynamics of an RLV-shaped, un-piloted, un-powered aircraft. The latter is a property wider than simple stability in that it allows vehicle instability to be accepted if the Stability Augmentation System can still enforce the desired dynamics. The proposed approach basically reduces the problem of determining the dynamic characteristics of the complete nonlinear system to the analysis of the robust stability of linear systems subject to uncertain parameters, by means of system’s linearization around a predetermined set of flight conditions. Then, the admissible uncertainties region is identified by means of a numerical code, based on a polynomial approach deriving from recent theoretical results for polynomials with uncertain coefficients. An application case on lateral-directional aerodynamic stability derivatives for the CIRA USV-FTB1 autonomous RLV demonstrator vehicle is analyzed. Results show the method’s ability to identify the maximum admissible uncertainties, and to address the areas of major concern. Since the admissible uncertainties region is a five dimensional set, a geometric projection-based visualization tool has been developed to provide meaningful insight into the admissible uncertainty region shape and extension.

This paper focuses on the autonomous real-time relative navigation of LEO satellite formations. Specifically, a novel closed loop approach which integrates an Extended Kalman Filter with an Integer Least Squares estimator is presented in which double differenced code and carrier measurements on two frequencies are processed to get accurate relative positioning. Real-world GPS measurements from the GRACE mission are used for assessing the positioning algorithm performance. Results demonstrate that the approach is suitable for real-time relative positioning with a centimeter-level accuracy.